1. Field of the Invention
This invention relates in general to a method of manufacturing composites and in particular to a method for manufacturing composite panels and stiffeners by first preparing a circumferential array of material and then forming and curing the material to the desired shape.
2. Description of Related Art
Composite structures and parts are necessary parts for industries requiring high strength, lightweight materials. A good example of this is in the aerospace industry, where aircraft and other airborne vehicles require high strength components that weigh as little as possible.
Many approaches have been previously developed for forming multiple layers of composite material into a desired shape or shapes. The most common, particularly in the aircraft industry, involves placing individual layers of material onto a form having a desired shape, and then curing the layers. Curing the material through application of heat and pressure fully compacts or debulks the composite material. The cured composite material then has the desired shape and strength. Forming parts in this way does not involve significant reshaping of the composite material during curing and may be very time consuming.
Disadvantages inherent in the aforementioned process include the very tedious and time consuming operation of laying individual layers of composite material directly onto a tool to obtain a final non-flat desired shape. The very labor intensive process of placing the layers of material onto a form may require many highly-skilled man-hours for each part, and is, therefore, very expensive. Additionally, the aforementioned process may require stopping after placement of every few layers of material and providing some form of mechanical compaction to the material. This may be necessary to achieve final full compaction of the layers. Failure to achieve full compaction of the material layers prior to curing may result in wrinkles and other anomalies in the final structure, since as individual layers compact, the local path-lengths of the fibers in the layers change. Wrinkles and other anomalies in the cured structure are aesthetically and structurally undesirable.
Previously developed methods for forming composite parts also fail to assure uniformity between parts. In many prior art methods, each part is separately made. Each part is formed by the process of placing individual layers onto a form and then curing the layers while on the form. The cured part is removed from the form, allowing the next part to be made by the same process. By this method a number of parts can be formed. Unfortunately, variations in compaction, in resin bleeding from the part, and in fiber xe2x80x9cwashingxe2x80x9d or dislocations from resin bleeding, tend to occur because compaction is occurring three dimensionally, and because of the low viscosity of the resin. These factors may yield parts that lack uniformity. Previously developed methods for building composite parts are, therefore, not compatible with low-cost, high-volume manufacturing methodologies.
Composite parts fabricated by previously developed methods often require machining after curing, e.g. routing, grinding, etc., in order to meet final dimension requirements. This machining adds additional time and expense to the process of fabricating the part and can result in damage to the part by delamination of the cured layers.
Yet another disadvantage of the previously developed methods for fabricating composite parts is their incompatibility with in-process control (IPC), statistical process control (SPC), and total quality (TQ) methodologies. IPC, SPC, and TQ require repeatable, measurable results to obtain full effectiveness. The custom approach of the prior art to fabricating composite parts is not amenable to obtaining the benefits of IPC, SPC, and TQ, i.e., high quality, high yield, and low cost.
Previously developed methods for forming composite parts often do not provide acceptable results when forming complex parts from two or more sub-parts or pre-forms by xe2x80x9cco-curingxe2x80x9d. In co-curing, two or more sub-parts are made into a single part by placing the sub-parts in the desired orientation and curing the combination. Since the prior art requires the individual layers of a sub-part to be laid-up in their final shape on a form joining two or more individual sub-parts to make a part, e.g., two channels and two plates to form an I-beam, is very difficult. If a foreign material, e.g.,backing paper or tape, is accidentally trapped between the layers during layup of a part, there is little likelihood that it will be detected. As a result, high labor costs may be invested in a complex, co-cured part that must be scrapped due to the inclusion.
Prior methods for fabricating composite parts often require that the individual composite layers be stored in a freezer prior to layup. This adds additional handling and equipment costs in fabricating a composite part.
Therefore a need has arisen for an improved method and system for fabricating parts from composite materials.
A need further exists for an improved method and system for reducing the time necessary for fabricating composite parts.
A further need exists for a low-cost method and system for fabricating parts from composite materials.
Yet another need exists for a method and system for fabricating multiple uniform parts from composite materials that do not require significant amounts of machining after curing.
Another need exists for a method for fabricating composite parts compatible with IPC, SPC, and TQ.
Yet another need exists for a method and system for fabricating composite parts that eliminate the need for special handling and storage of composite layers.
The present invention provides a method for manufacturing a composite part having a specified shape, thickness, and density. The composite part is formed in an array of similar parts located around a mandrel having an outer surface and a longitudinal axis. Several tooling members are disposed around the outer surface of the mandrel, the tooling members serving as the molds around which the composite parts will be formed. A plurality of filler members are disposed on the outer surface of the mandrel between the tooling members. An outer surface of each tooling member and an outer surface of each filler member combine to form a generally smooth and rounded application surface that surrounds the mandrel.
A plurality of composite layers is positioned on the application surface using an automated positioning technique such as fiber placement. A part formation aid is placed on the composite layers above each tooling member, and a cut is made in the composite layers parallel to the longitudinal axis of the mandrel and between each tooling member. The filler members are then removed from between the tooling members.
Finally, the mandrel and composite layers are placed in a vacuum bag, which is then placed in an autoclave. The vacuum bag is an elastomeric material in the form of a tube that is slid over the mandrel. The vacuum bag tube can be pulled out of the way during fiber placement and slid over the mandrel at the time of forming. As the bag is evacuated, each part formation aid deforms toward the mandrel, thereby forming the composite layers around the tooling members. The composite part resulting from the evacuation of the bag is then cured for a specified amount of time to insure that the composite part will maintain its specified shape, thickness, and density.
Alternatively, the method according to the present invention is used to fabricate composite parts having more than one component. One application of this method is to produce a flanged panel having a plurality of hat-shaped stiffeners integrally disposed on the panel.
The flanged panel is produced using a rectangular mandrel having an outer surface. A filler member having an outer surface is placed on one end of the mandrel, the outer surface of the filler member and the outer surface of the mandrel forming a generally smooth and rounded application surface.
A plurality of composite layers is positioned on the application surface using an automated positioning method such as fiber placement. After placement of the layers, the filler member is removed, and the composite layers are subjected to the vacuum bagging technique previously mentioned. This process forms a flanged composite panel. The panel is then partially cured.
After partial curing, a plurality of tooling members are disposed on an outer surface of the recently formed panel. Filler members are placed on the outer surface of the panel between the tooling members to form a generally smooth second application surface. Composite layers are applied to the second application surface using fiber placement. The filler members are then removed and the composite layers are cut between the tooling members.
The flanged panel, the tooling members, and the newly applied composite layers are placed in a vacuum bag. As the bag is evacuated, the composite layers form into hat-shaped stiffeners around the tooling members. The flanged panels and the stiffeners are finally co-cured or co-bonded to form an integral composite part.